Exemplary embodiments pertain to the art of gas turbine engines, and more particularly to cooling of components of gas turbine engines.
Advancements in performance of turbomachines, such as gas turbine engines, performance has often been linked to overall pressure ratio of the turbomachine and a turbine inlet temperature that can be sustained during operation of the turbomachine. Increases in efficiency through increases in pressure ratio and/or turbine inlet temperature typically results in an increase in operating temperatures of turbine flow path components, in which temperatures of the working fluid in the turbine flow path is often several hundred degrees Fahrenheit higher than the melting point of component materials.
Components such as turbine vanes and blades and blade outer air seals, in the turbine section of the gas turbine engine are configured for use within particular temperature ranges. Often, the conditions in which the components are operated exceed a maximum useful temperature of the material of which the components are formed. Thus, such components often rely on cooling airflow to cool the components during operation. For example, stationary turbine vanes often have internal passages for cooling airflow to flow through, and additionally may have openings in an outer surface of the vane for cooling airflow to exit the interior of the vane structure and form a cooling film of air over the outer surface to provide the necessary thermal conditioning. Similar internal cooling passages are often included in other components, such as the aforementioned turbine blades and blade outer air seals.
Passages of various configurations have been used traditionally to cool turbine components. In the ongoing efforts to improve engine performance and efficiency, these configurations are becoming increasingly inadequate to provide sufficient cooling for the constituent materials. In these applications, dual-wall cooling may be utilized. These dual-wall passages are formed by thin ‘skin cores’ which provide a narrow cavity in the thickness direction between the main body core passage and the external hot gaspath wall. These passages may extend in any direction along the surface of the hot section component, which may be a blade, vane, outer air seal, combustor panel, or any other cooled component.